This invention relates generally to turbine engines and, more particularly, to turbine interface seals.
Gas turbine engines typically include a multistage axial flow low pressure compressor and a multistage axial flow high pressure compressor which supplies high pressure air to a combustor. The compressors include stages of stationary components referred to as stators and stages of rotational components, which add work to the system, referred to as rotors.
A portion of compressed high pressure air supplied to the combustor is mixed with fuel, ignited, and utilized to generate hot combustion gases which flow further downstream to one of the multistage flowpaths. Particularly, the combustion gases flow through one or more turbine stages which extract energy from the hot gases to power the rotors in the compressors and provide other useful work.
One turbine stage downstream from the combustor is commonly referred to as a turbine nozzle stage and includes a plurality of circumferentially spaced vanes that extend in a radial direction with respect to a central axis of the turbine engine. The vanes extend between an outer band and an inner band that assist in maintaining axial and radial positioning of the vanes and define a flowpath for the combustion gases.
Downstream from the nozzle stage is another rotor stage that includes a plurality of circumferentially spaced rotor blades that extend radially outward from a rotor disk and are surrounded by an annular shroud which also defines a flowpath for the combustion gases. Adjacent ends of the turbine nozzle outer band and rotor shroud are spaced apart by a gap to facilitate assembly as well as to accommodate differential thermal expansion and contraction that occurs during operation of the engine. The gap, however, also is a potential leakage path for compressed air.
Particularly, a portion of the compressed air may be extracted from the high pressure compressor for turbine section cooling, airframe pressurization, anti-icing, and other uses. For example, for turbine section cooling, a portion of the extracted air is channeled through the nozzle vanes. The extracted air is at a higher pressure than the combustion gases and is channeled to a cavity formed by a portion of a stator casing, an inside face of the outer band, and a portion of a rotor shroud hanger. The extracted air naturally seeks to move from the cavity to the combustion gas flowpath formed by an outside face of the outer band and the rotor shroud. Therefore, the gap between the outer band and shroud must be properly sealed, otherwise the high pressure extracted air would leak through the gap, or an interface, to the lower pressure flowpath.
It is known to use a "w" seal or a leaf seal to seal the interfaces in turbine engines. For example, in a typical turbine, outer bands extend substantially parallel to a central axis of the engine and include an axially spaced apart forward rail and aft rail. The rails extend radially outward from the outer bands with the aft rail of the nozzle outer band adjoining an adjacent rotor shroud or rotor shroud hanger. Typically, the leaf seals are assembled and mounted to the aft rails with mounting pins. The mounting pins extend through an inside face of the aft rails.
Sealing the interfaces in turbine engines and, in particular, a turbine nozzle is difficult due to the location of the nozzle, the increases and decreases in the temperature of the turbine engine components, and the resultant increases in the temperature and degradation of the seals. It would be desirable to provide a turbine engine interface seal that is easy to install and can withstand high temperatures for long periods of time and retain the necessary sealing characteristics.